Gas turbine combustor, particularly for an aircraft engine

ABSTRACT

A gas turbine combustor, in particular for an aircraft engine, has at least one chamber, through which combustion gas flows in use, and which is defined by a lateral wall with channeling for feeding cooling air into the chamber and cooling the lateral wall; part of the channeling is defined by at least one double wall for guiding a relative stream of cooling air into the chamber in a tangential direction with respect to the lateral wall.

[0001] The present invention relates to a gas turbine combustor,particularly for an aircraft engine.

BACKGROUND OF THE INVENTION

[0002] As is known, jet aircraft engines comprise a compressor; anexpansion turbine; and a combustor interposed between the compressor andthe turbine. The combustor comprises a combustion chamber communicatingwith the compressor outlet; and a turbine inlet chamber or conduit,along which, in use, flows relatively high-temperature gas generatedinside the combustion chamber.

[0003] To reduce thermal stress caused by the high gas temperature, theinner walls defining the turbine inlet conduit and the combustionchamber must be cooled continually; for which purpose, the inner wallshave a number of through holes, through which relatively low-temperatureair flows crosswise to the inner walls into the conduit where it mixesdirectly with the gas flowing towards the turbine.

[0004] Known combustors of the above type are unsatisfactory byrequiring a relatively large amount of air to cool the inner walls asrequired, and which is mainly due to inefficient heat exchange betweenthe inner walls and the airflow through the holes formed in the innerwalls.

[0005] Moreover, the holes produce stress concentrations in the materialin which they are formed, and so tend to impair the mechanical strengthof the inner walls of the combustor.

SUMMARY OF THE INVENTION

[0006] It is an object of the present invention to provide a gas turbinecombustor, particularly for an aircraft engine, designed to provide astraightforward, low-cost solution to the above problems, and which inparticular provides for improving heat exchange between the cooling airand the inner walls.

[0007] According to the present invention, there is provided a gasturbine combustor, in particular for an aircraft engine; the combustorcomprising at least one chamber, through which combustion gas flows inuse; at least one lateral wall defining said chamber; and channelingmeans associated with said lateral wall to permit the passage of acooling fluid for cooling the lateral wall; and being characterized byalso comprising guide means at least partly defining said channelingmeans to feed at least one stream of said cooling fluid into saidchamber in a tangential direction with respect to said lateral wall.

BRIEF DESCRIPTION OF THE DRAWINGS

[0008] A non-limiting embodiment of the invention will be described byway of example with reference to the accompanying drawings, in which:

[0009]FIG. 1 shows a schematic, partial diametrical section of apreferred embodiment of the gas turbine combustor, particularly for anaircraft engine, according to the present invention;

[0010]FIG. 2 shows a larger-scale, diametrical half-section of the FIG.1 combustor.

DETAILED DESCRIPTION OF THE INVENTION

[0011] Number 1 in FIG. 1 indicates as a whole a jet aircraft engine,which is axially symmetrical with respect to an axis 3, and comprises acompressor (not shown), a combustor 5 (shown partly), and a turbine 6(shown partly and schematically), arranged in series with one anotheralong a gas-flow path 7 through engine 1.

[0012] With particular reference to FIG. 2, combustor 5 comprises asupporting structure 8 (shown partly), and defines an inner annularcavity 9 having an inlet (not shown) communicating with the outlet ofthe compressor, and an outlet 10 communicating with the inlet of turbine6. Cavity 9 comprises a combustion chamber 11 (shown partly) defined bytwo facing walls 12 and 13; and an annular chamber 14 formed in anintermediate position between chamber 11 and outlet 10 to feed the gasinto turbine 6.

[0013] Chamber 14 decreases gradually in section towards outlet 10 toaccelerate flow of the gas, and is defined by two facing walls 15 and 16having a curved diametrical section and converging with each othertowards outlet 10. Wall 16 is convex towards chamber 14, forms anextension of wall 13, and has a number of holes 17 by which streams ofcooling air flow through wall 16.

[0014] Wall 15, on the other hand, is defined by an annular structurecomprising a fastening end portion 21 (shown partly), which extendsoutwards of cavity 9 and is connected, on one side, to structure 8 inknown manner not shown in detail, and, on the other side, to wall 12 bymeans of an annular flange 23 integral with wall 12. Portion 21 isconnected to structure 8 and to flange 23 to allow structure 15 arelatively small amount of axial and radial movement to compensate, inuse, for high-temperature-gradient deformation.

[0015] Again to compensate for high-temperature-gradient deformation,structure 15 also comprises an annular end portion 25 facing outlet 10and connected to structure 8 to slide along a curved guide 26 (shownpartly) substantially parallel to path 7.

[0016] More specifically, portion 25 comprises a wall 27, which definesand is concave towards chamber 14, and which has a number of holes 28for the passage of cooling air streams through wall 27.

[0017] Structure 15 also comprises an intermediate portion 30 betweenportions 21 and 25, and comprising two annular walls 31 and 32, whichextend facing each other and spaced apart along path 7, with theirrespective concavities facing chamber 14. Wall 31 is a seamlessextension of wall 27, whereas wall 32 defines chamber 14 and comprisestwo opposite panels 34 and 35 respectively facing wall 31 and an endportion 36 of wall 12. Panels 34 and 35 project from an intermediatefastening portion 38, which is integral with portion 21 and U-shapedwith its concavity facing panel 34.

[0018] Panel 35 and portion 36 define an annular opening 40 formed in anideal surface perpendicular to path 7; and an annular guide channel 41,which is substantially parallel to path 7, communicates with a coolingair inlet 43 formed in flange 23, and comes out inside chamber 14through opening 40.

[0019] Panel 34 and wall 31, on the other hand, converge with each othertowards outlet 10, and define an annular opening 45 formed in an idealsurface perpendicular to path 7; and an annular guide channel 47, whichis substantially parallel to path 7, communicates with a cooling airinlet 48 formed in portion 38, and comes out inside chamber 14 throughopening 45.

[0020] More specifically, panel 34 has a substantially smooth surface 50defining chamber 14; and a surface 51 defining channel 47 and havingcircumferential ribs 52.

[0021] With reference to FIG. 2, portion 30 comprises a number of stopmembers carried by wall 31, angularly spaced about axis 3, and only oneof which is shown and indicated 55 in FIG. 2. Member 55 projects insidechannel 47, is detached from surface 51, and defines a stop for the freeend 56 of panel 34 when free end 56 moves towards wall 31.

[0022] In actual use, panels 34, 35 and portion 36 of wall 12 definechannels 41 and 47, and guide two cooling air streams F1 and F2 intochamber 14 in respective directions tangential with respect to structure15, and, more specifically, in directions concordant with each other andwith gas-flow path 7. As they flow along structure 15, streams F1, F2cool structure 15 and gradually mix with the combustion gas flowing inchamber 14. More specifically, stream F1 flows tangentially with respectto wall 32, while stream F2 is accelerated along channel 47 by thegradually narrowing section of channel 47, flows into chamber 14tangentially with respect to wall 31, and flows along wall 27 togetherwith stream F1.

[0023] At the same time, free end 56 is movable crosswise to wall 31 andpath 7 to vary the flow section of opening 45 and channel 47 as afunction of the structure 15 determined by the gas temperature, and ismoved by the deformation of panel 34 and U-shaped portion 38 caused bythe temperature gradient of various operating conditions. Morespecifically, as temperature increases, panel 34 moves automaticallytowards wall 31, so that the section of opening 45 and channel 47narrows to increase the speed of stream F2.

[0024] The increase in the speed of stream F2 increases the amount ofheat removed from, and so reduces the temperature of, structure 15, sothat free end 56 moves away from wall 31 to reduce the speed of streamF2 and, therefore, the amount of heat removed. In the steady operatingcondition of combustor 5, the position of panel 34 settles after anumber of cycles, so that the flow sections of opening 45 and channel 41reach a balance condition.

[0025] In combustor 5 described and illustrated, channels 41, 47 andopenings 40, 45 therefore form part of a channeling system 60, whichprovides for cooling structure 15 with greater heat-exchange efficiencythan by feeding cooling air through holes formed crosswise to structureor wall 15.

[0026] Streams F1, F2, in fact, are guided tangentially with respect tothe surfaces defining chamber 14, so that heat exchange between thecooling air and wall or structure 15 takes place over a relatively widearea.

[0027] Consequently, a much smaller amount of air is needed to keepstructure 15 below a given required temperature: in particular, roughlyhalf the amount required using a structure 15 simply provided withthrough holes.

[0028] Moreover, heat-exchange efficiency is also improved by providinga first and second stream F1, F2 flowing successively in concordantdirections, and the effects of which are combined at the point in whichcooling by stream F1 tends to become less effective. The way in whichthe two successive streams F1, F2 are provided is also extremelystraightforward, by one wall 32 defining both openings 40, 45.

[0029] Adjusting the speed of stream F2 by means of thermal deformationof panel 34 and portion 38 also provides for automatically regulatingcooling of structure 15 as a function of temperature, and is achievedextremely easily by panel 34 projecting from portion 38.

[0030] Members 55 also provide for controlling the flow section ofstream F2, in the sense of preventing full closure of opening 45.

[0031] Structure 15 also has good mechanical structural characteristics,by having substantially no through holes which would tend to weaken it.

[0032] Clearly, changes may be made to combustor 5 as described hereinwithout, however, departing from the scope of the present invention.

[0033] In particular, openings 40, 45 and channels 41, 47 may differfrom those described and illustrated, e.g. to direct streams F1, F2 inslightly different tangential directions, and/or with a circumferentialcomponent with respect to axis 3; and openings 40, 45, as opposed tobeing annular, may be formed along only part of the circumferentialperiphery of cavity 9.

[0034] The cooling fluid may be other than air; and/or more than twosuccessive tangential streams may be provided along the periphery ofchamber 14.

[0035] Finally, the present invention may also be applied to walls 16,27 defining chamber 14, and/or to walls 12, 13 defining combustionchamber 11.

1) A gas turbine combustor (5), in particular for an aircraft engine(1); the combustor comprising at least one chamber (14), through whichcombustion gas flows in use; at least one lateral wall (15) definingsaid chamber (14); and channeling means (60) associated with saidlateral wall (15) to permit the passage of a cooling fluid for coolingthe lateral wall (15); and being characterized by also comprising guidemeans (31, 32, 36) at least partly defining said channeling means (60)to feed at least one stream (F1, F2) of said cooling fluid into saidchamber (14) in a tangential direction with respect to said lateral wall(15). 2) A combustor as claimed in claim 1, characterized in that saidguide means (31, 32, 36) comprise at least one pair of walls (31, 32)forming part of said lateral wall (15). 3) A combustor as claimed inclaim 1, characterized in that said guide means (31, 32, 36) define atleast one channel (47) decreasing in section towards an inlet (45) intosaid chamber (14). 4) A combustor as claimed in claim 1, characterizedin that said guide means (31, 32, 36) define a first and a second inlet(40) (45) formed in respective spaced portions to feed respectivetangential, substantially concordant streams (F1) (F2) of cooling fluidinto said chamber (14). 5) A combustor as claimed in claim 4,characterized in that said first and said second inlet (40) (45) areformed in succession along a path (7) of said gas in said chamber (14).6) A combustor as claimed in claim 4, characterized in that said guidemeans (31, 32, 36) comprise first (31) (35) and second (34) (36) guidemeans for each said first and second inlet (40) (45); said first guidemeans (35) associated with said first inlet (40), and said second guidemeans (34) associated with said second inlet (45) comprising a commonwall (32). 7) A combustor as claimed in claim 1, characterized bycomprising speed adjusting means (34, 38) for adjusting the speed ofsaid stream (F2) through said inlet (45). 8) A combustor as claimed inclaim 7, characterized in that said speed adjusting means (34, 38)comprise section adjusting means (34, 38) for reducing the section ofsaid inlet (45) automatically alongside an increase in the temperatureof said lateral wall (15). 9) A combustor as claimed in claim 8,characterized in that said section adjusting means comprise at least onedeformable member (34, 38) deformable as a function of said temperature.10) A combustor as claimed in claim 2, characterized in that one of saidwalls (32) comprises a fastening portion (38), and a panel (34) partlydefining said inlet (45); said deformable member (34, 38) being definedby at least one of said panel (34) and said fastening portion (38). 11)A combustor as claimed in claim 10, characterized in that said panel(34) projects from said fastening portion (38). 12) A combustor asclaimed in claim 11, characterized in that said fastening portion (38)is U-shaped with its concavity facing said panel (34). 13) A combustoras claimed in claim 10, characterized in that said panel (34) has asubstantially smooth surface (50) defining said chamber (14); and aribbed surface (51) defining said inlet (45). 14) A combustor as claimedin claim 1, characterized in that each said inlet (40, 45) is defined bya respective annular opening about an axis (3) of said combustor (5).